100results about How to "Improve flight performance" patented technology

The invention discloses a helicopter mountain flight simulation and flight strategy design method and related devices, and relates to the field of aerodynamics and computational fluid mechanics of rotorcrafts, the method comprises the following steps: establishing a mountain terrain model and a mountain wind environment model and carrying out numerical simulation to obtain flow field velocity distribution and turbulence energy distribution of a whole watershed; a real mountain environment is provided for follow-up helicopter mountain flight simulation, interference and response characteristics encountered by a helicopter in mountain flight can be predicted more accurately, a mountain wind environment model can simulate fluctuating wind and average wind, and the inflow velocity of a whole-basin flow field changing along with time and height is provided; flight strategy research in a complex mountain environment is carried out through rotor aerodynamic loads, full-aircraft response characteristics and driver control loads, a helicopter flight simulation technology based on environment-helicopter-pilot multi-factor coupling is formed, and the flight performance and safety of a helicopter in the complex environment are improved.

The invention relates to the technical field of database establishment, in particular to an establishment method and system based on an aircraft aerodynamic model database. The method comprises the following steps: obtaining an aircraft pneumatic data storage demand and an aircraft pneumatic data query demand, and carrying out multi-level storage architecture design to obtain a basic data layer, a processing data layer and a high-level model data layer; the method comprises the following steps: collecting aircraft original aerodynamic data sets of wind tunnel experiments, flight tests and CFD numerical simulation, and transmitting and storing the aircraft original aerodynamic data sets to a basic data layer; carrying out aerodynamic load, flow field parameter and aerodynamic pressure distribution analysis on the original aerodynamic data set of the aircraft in the basic data layer, and transmitting and storing the aerodynamic load, flow field parameter and aerodynamic pressure distribution to the processing data layer; and carrying out aircraft aerodynamic model construction on the data in the processing data layer, transmitting and storing the data to a high-level model data layer, and carrying out R-tree multi-dimensional index connection to obtain an aircraft aerodynamic model multi-level storage database. According to the invention, an efficient and intelligent aerodynamic model database can be established.

The invention discloses an aircraft folding type propeller mechanism and an aircraft, the aircraft folding type propeller mechanism comprises a rotating shaft, a driving part, a first propeller and a second propeller; the driving piece is connected with one end of the rotating shaft; the first propeller comprises a first propeller hub and a first propeller blade which are connected with each other, the first propeller hub is fixedly connected to the rotating shaft, a thrust groove is formed in the side, away from the driving part, of the first propeller hub, and a first thrust surface which is obliquely arranged is arranged at the bottom of the thrust groove; the second propeller comprises a second propeller hub and a second propeller blade which are connected, the second propeller hub is movably arranged on the rotating shaft in a sleeving mode and located on the side, away from the driving piece, of the first propeller hub, and a thrust block extending into the thrust groove is arranged on the side, facing the first propeller hub, of the second propeller hub in a protruding mode; and the thrust blocks are matched with the thrust grooves, so that the second paddles are in a folded state or an unfolded state relative to the first paddles. According to the folding propeller mechanism of the aircraft, the resistance of the aircraft in the flight process is reduced, and the flight performance of the aircraft is improved.

The invention discloses a rotor aerodynamic force and flow field simulation method, and relates to the field of aerospace, and the method comprises the steps: generating a blade skin grid according to rotor configuration parameters, and constructing a Cartesian background grid; constructing a calculation model of variable-diameter rotor flow field simulation based on the blade skin grid and the Cartesian background grid; and in the constructed calculation model for variable-diameter rotor flow field simulation, a high-precision CFD method suitable for variable-diameter rotor flow field solving is established, rotor aerodynamic force and a flow field in the dynamic diameter changing process are simulated, and a CFD calculation result is output. The aerodynamic performance of the rotor wing can be improved by simulating the variable-diameter rotor wing flow field.

The invention belongs to the field of trajectory optimization, and provides an unmanned aerial vehicle-unmanned vehicle joint trajectory optimization method and system, a medium and equipment, and the technical scheme is that the trajectory characteristics of an unmanned aerial vehicle and an unmanned vehicle are integrated, and the motion characteristics, environmental constraints and task requirements of the unmanned aerial vehicle and the unmanned vehicle are comprehensively considered. The cooperative task of the unmanned aerial vehicle and the unmanned vehicle is completed in a high-efficiency and high-quality manner, it is ensured that the unmanned aerial vehicle and the unmanned vehicle fully excavate the cooperative ability, unconstrained trajectory optimization is adopted, all constraint penalty terms are set as soft constraint terms, the trajectory is specifically required according to a specific task, and if the trajectory is required to pass a specific point, the unmanned aerial vehicle and the unmanned aerial vehicle are matched with each other. New constraint penalty terms can be conveniently added to the trajectory; adjustment of cooperation of an energy consumption constraint penalty term and a time regular term is added, so that the energy consumption of the unmanned aerial vehicle is reduced; landing terminal constraint is carried out on the trajectory with the landing task, so that the unmanned aerial vehicle obtains safe and efficient landing quality at the confluence point.

The invention discloses an unmanned aerial vehicle pitching control method and system for optimizing PID through fuzzy immune control, and relates to the field of unmanned aerial vehicle control, and the analysis method comprises the steps: carrying out the decoupling and linear simplification of a motion equation of an unmanned aerial vehicle, obtaining a longitudinal motion mathematical model, carrying out the parameter initialization of a PID controller according to the response requirements and performance indexes of the system, and obtaining a longitudinal motion mathematical model; the method comprises the following steps: establishing an unmanned aerial vehicle PID control system, establishing a fuzzy control module, carrying out fuzzy reasoning adjustment, initializing immune algorithm parameters, determining an optimization termination condition of an immune algorithm, obtaining an initial unmanned aerial vehicle PID control system, optimizing the initial unmanned aerial vehicle PID control system, connecting the optimized unmanned aerial vehicle PID control system with a controlled object, and operating the optimized unmanned aerial vehicle PID control system. According to the method, the purpose of adaptively adjusting the PID parameters can be achieved, different flight environments and task requirements are adapted, and the flight performance and task execution efficiency of the unmanned aerial vehicle are improved.

The invention relates to the technical field of aircraft control, in particular to an airfoil gust retarding feedforward control method based on a circular rector control technology, which comprises the following steps: modifying the shape of the rear edge of an NACA0012 airfoil, and arranging jet holes in the rear edge to form a circular rector control airfoil; performing a fluid simulation experiment on the circular rector control airfoil based on the gust speed, and constructing a model of the airfoil lift coefficient disturbed by the gust speed; performing a fluid simulation experiment on the circular rector control airfoil based on the jet flow momentum coefficient, and constructing a model of which the airfoil lift coefficient is controlled by the jet flow momentum coefficient; based on the model, obtaining a gust slowing-down feedforward control model of which the input quantity is the vertical gust speed and the output quantity is the jet flow momentum coefficient; controlling jet flow of a trailing edge by adopting the gust slowing-down feedforward control model to complete gust slowing-down control of the circular rector control airfoil; the method can improve the ability of the aircraft to resist the gust disturbance environment.

The ornithopter comprises a main beam, folding and unfolding steering engines are symmetrically installed on the two sides of the main beam, and humerus assemblies and second connecting rods located below the humerus assemblies are arranged on the two sides of the main beam; a radius component is movably connected to one end, far away from the main beam, of the humerus component; one end of the second connecting rod away from the girder is movably connected with a first connecting rod; the radius assembly is movably connected with the first connecting rod through the wing rib assembly, a parallelogram-shaped folding and unfolding structure is formed, the folding and unfolding steering engine is connected to the second connecting rod through the driving assembly, and the folding and unfolding steering engine can drive the folding and unfolding structure to fold and unfold. The wing ribs and the radius form a parallelogram folding and unfolding structure, the strength of the radius is enhanced, the whole folding and unfolding structure is movably connected, the folding and unfolding structure is driven to move smoothly and is not prone to being blocked, the driving stroke of the steering engines is short, folding and unfolding response is fast, the wings on the two sides are independently controlled by the steering engines to be folded and unfolded, and rapid steering of the ornithopter can be achieved. And the flying capability of the folding and unfolding wings is improved.

The invention discloses a flight performance improvement method based on aero-engine emergency control, and the method comprises the steps: calculating a flight envelope of an aircraft in an original limit state according to the parameters of the aircraft; according to the flight envelope in the original limit state, an engine thrust increasing controller is designed to carry out emergency control on the engine, and the flight performance is improved; and calculating the flight envelope of the aircraft after the flight performance is improved, comparing the flight envelope with the flight envelope in the original limit state, and carrying out quantitative evaluation on the flight performance improvement. By adjusting the control strategy of the aero-engine, higher engine thrust can be obtained by sacrificing part of the service life of the engine in emergency. The flight performance of the aircraft under the emergency condition can be effectively improved, the coping capacity and safety of the aircraft under the emergency condition can be improved, the application range of the aircraft in the complex environment is expanded, and greater safety guarantee is provided for pilots and passengers.

The invention provides a method and a system for measuring and calculating lateral deviation of a landing runway, and relates to the technical field of civil aviation safety. The method comprises the following steps: acquiring a corresponding relation between a runway distance and a deviation angle and course track deviation according to a landing runway design specification; the runway distance is the distance from the runway entrance to the course beacon station; obtaining a runway entrance mark line, and reversely circulating a flight motion track according to the flight QAR data and the flight landing point; obtaining a runway entrance height based on an intersection point generated by the flight motion trail and the runway entrance mark line; and obtaining a course track deviation parameter corresponding to the runway entrance height, and calculating a lateral deviation distance based on the runway distance and the corresponding relation. The position of the airplane reaching the runway entrance is recognized based on landing point reverse circulation, so that accurate transverse deviation is solved according to the course track deviation of the entrance position and the runway distance, and the accuracy of a calculation result is improved; meanwhile, related parameters are only based on landing runway design specifications, aircraft type information does not need to be associated, and universality is higher.

The utility model relates to an unmanned aerial vehicle air-drop simulation training bomb, which comprises a touch switch, a fuse, a warhead and an empennage, and is characterized in that the warhead is filled with explosives; the touch switch is fixedly arranged at one end of the fuze, the end, away from the touch switch, of the fuze is fixedly connected with the warhead, and the end, away from the fuze, of the warhead is fixedly connected with the empennage; the touch switch is coupled with a time delay circuit for delaying the trigger fuze; the fuse is provided with an electric detonator interface, and the electric detonator interface is coupled with the time delay circuit. According to the invention, trainees can better master the specific time efficiency of unmanned aerial vehicle air-drop ammunition percussion detonation in training, and the training effect is improved.

The invention provides a control and unfolding locking system suitable for an n-shaped folding wing of a pipe-launched unmanned aerial vehicle, and belongs to the technical field of pipe-launched unmanned aerial vehicle control. The system comprises a rotary driving module for driving the main wing and the outer wings on the two sides to rotate and unfold from the folding position and an unfolding locking module for locking the main wing and the outer wings on the two sides to the complete unfolding position, and is further provided with an outer wing folding locking module for locking the outer wings on the two sides to the folding position and a control module for controlling the outer wings to be unlocked in a time-sharing mode. The wing on one side rotating forwards unlocks the outer wing when the main wing rotates to be perpendicular to the fuselage, and the wing on one side rotating backwards unlocks the outer wing before the other outer wing for a preset time, so that a rolling restoring moment is generated to balance a rolling moment generated by the difference of incoming flow speeds on the wing surfaces of the wings on the two sides at the initial stage of unfolding of the main wing; the problem that the rolling posture of the unmanned aerial vehicle is large in deviation after the unmanned aerial vehicle leaves the tube due to the control mode that the outer wing is restrained and folded in the unmanned aerial vehicle launching tube through the fuselage and is automatically unfolded after the unmanned aerial vehicle is launched and leaves the tube is solved.

The invention discloses a flutter control method and a flutter control device for a high-aspect-ratio wing. The flutter control method comprises the following steps: acquiring an initial flutter speed of the wing; steel cables are installed on the wings, and pre-tightening force is applied to the steel cables in the flight state through a servo mechanism so as to limit torsion of the wings; acquiring the expansion flutter speed of the wing after the pre-tightening force is applied to the steel rope; starting a wingtip angular displacement sensor in the flight process of the unmanned aerial vehicle, wherein the wingtip angular displacement sensor is used for monitoring a wingtip torsion angle and a torsion angular speed in real time; and when the flight speed of the unmanned aerial vehicle exceeds the extended flutter speed, a servo mechanism is used for dynamic control. The problem of how to improve the flutter speed and enable the unmanned aerial vehicle to safely fly at a higher speed so as to improve the flight performance of the unmanned aerial vehicle is solved.

The invention provides an aircraft equal-straight-part skin assembly and an aircraft equal-straight-part framework, the aircraft equal-straight-part skin assembly comprises a plurality of wall plates which are annularly and sequentially spliced, the inner surfaces of the wall plates are flush, and the inner surfaces of all the wall plates form a smooth cylindrical surface; the outer surface of each wall plate comprises a plurality of equal-thickness areas and transition areas which are alternately arranged in the heading direction, the thickness of any equal-thickness area is not smaller than the corresponding minimum load thickness, the surface of any transition area is an annular face arranged in the annular direction, and the boundaries of the two sides of any transition area make contact with the adjacent equal-thickness areas correspondingly. The thicknesses of the boundaries on the two sides of the transition areas are equal to the thicknesses of the equal-thickness areas in contact with the transition areas, and any transition area has a preset gradient in the course; therefore, under the condition that the aerodynamic performance of the aircraft is guaranteed, the configuration of the stringer and the annular frame tightly attached to the inner surface of the aircraft skin is simplified, and the manufacturing cost of the aircraft is reduced.

The utility model discloses an integrally-formed unmanned aerial vehicle carbon fiber arm which comprises an upper arm part, a reinforcing arm is fixed to the bottom end of the upper arm part, a lower arm part is fixed to the end, away from the upper arm part, of the reinforcing arm, and a second side connecting arm is fixed to the outer wall of one side of the upper arm part. A first side connecting arm is fixed to the outer wall of the side, away from the second side connecting arm, of the upper arm part, a part containing seat is arranged on the outer wall of the side, away from the lower arm part, of the first side connecting arm, and unmanned aerial vehicle seats are installed at the two ends of the second side connecting arm correspondingly. According to the unmanned aerial vehicle arm, the problems that a traditional unmanned aerial vehicle arm is large in weight, insufficient in mechanical strength and poor in fatigue resistance are solved, the overall weight of the unmanned aerial vehicle arm is greatly reduced, and the unmanned aerial vehicle arm can bear large static loads and high-frequency vibration loads without fracture deformation; and impact and vibration can be resisted to a certain extent, and the aim of easily assembling and connecting the unmanned aerial vehicle arms is achieved.

The invention relates to the field of electronic engineering, and provides a flying attitude regulation and control method and system based on a capacitive fixed wing, and the method comprises the steps: firstly obtaining driving dual motors of a fixed-wing unmanned aerial vehicle, collecting and preprocessing real-time motor data, extracting key parameters, then recognizing a connected adjustable resistor and an adjustable end, calculating the total resistance value, and calculating the resistance value of the fixed-wing unmanned aerial vehicle according to the resistance value of the fixed-wing unmanned aerial vehicle. And determining a triode connection mode, grounding to generate a current loop, and calculating a collector end current value. And then determining an influence factor vector based on the total resistance value and the current value, correcting the flight attitude, analyzing flight environment parameters and calculating a compensation value, constructing and optimizing a differential control strategy by combining the compensation value and a motor state, generating a control instruction, and monitoring feedback data after execution to generate an attitude regulation report. The flying performance and controllability of the capacitive fixed wing can be improved.

The invention relates to the technical field of aircrafts, and particularly discloses a fixed wing aircraft capable of vertically taking off and landing, which comprises a fixed wing, a storage box fixedly connected to the bottom of the fixed wing, a plurality of supporting legs connected to the outer wall of the storage box, and a rotating mechanism used for changing the flight direction of the aircraft, the rotating mechanism is positioned on the fixed wing; the lubricating mechanism is used for lubricating the rotating mechanism, and the lubricating mechanism is located on the rotating mechanism; through the action of the rotating mechanism and the action of controlling the rotor motors to work by using the attitude sensors, the connecting plates can be driven to rotate according to the lift force difference generated by the rotor motors, so that the aircraft can fly in a horizontal or vertical state, and the aircraft can be changed in multiple directions in the air; and the flight modes of the aircraft are increased, so that the flight use effect of the aircraft is improved.

The utility model discloses an aircraft recovery device and an aircraft, and relates to the technical field of aircrafts. The recycling bin is installed on the foundation, the end, away from the ground, of the recycling bin is of an opening structure, and a plurality of through holes are formed in the side wall of the recycling bin in the circumferential direction; the fixing mechanisms are arranged on the outer side of the recycling bin in the circumferential direction; the multiple locking mechanisms are rotationally connected to the multiple fixing mechanisms corresponding to the locking mechanisms, at least part of the locking mechanisms penetrate through the through holes and are located in the recycling barrel, and when the aircraft is recycled, the part of the locking mechanisms located in the recycling barrel abut against the bottom of the aircraft. And the locking mechanism located outside the recycling bin penetrates through the through hole and abuts against the outer wall of the aircraft so as to lock the aircraft. Therefore, according to the aircraft recovery device, the aircraft is prevented from carrying an undercarriage, the energy consumption of the aircraft is reduced, the flight performance of the aircraft is improved, and meanwhile, the recovery safety of the unmanned aerial vehicle is improved.

The utility model provides a freight unmanned plane relates to unmanned plane technical field, including fuselage for loading goods, the fuselage tail lower portion is equipped with rear aircraft wheel, rear aircraft wheel is connected with the fuselage through support rod, rear wing is located the fuselage tail upper portion and has upper reflex angle, front wing is located the fuselage head lower portion and has lower reflex angle, the front wing both sides lower portion is equipped with front aircraft wheel, tail push propeller is located the fuselage tail, vertical propeller is located the fuselage both sides, wherein, tail push propeller drives through the fuselage upper oil dynamic piston engine, vertical propeller drives through fuselage upper motor. A freight unmanned plane in the utility model, solved the problem that the prior art lacks a freight unmanned plane capable of stable vertical take-off and landing and long range and large load.

The invention provides a jet circulation empennage control system, an empennage and an aircraft, and relates to the technical field of active flow control, and the jet circulation empennage control system comprises a circulation controller, a steering engine control module, an auxiliary power system, an empennage circulation control unit, an airflow monitoring sensor and an airflow valve. Compressed air is provided by an auxiliary power system of the aircraft, so that the weight of the aircraft can be reduced while yaw control and / or pitching control are / is realized, the number of parts of the whole aircraft is reduced, and the reliability of the system is improved. The aircraft does not need a hydraulic system with a complex structure any more, the weight of the aircraft is reduced, the consumption of fuel oil is reduced, and therefore the economical efficiency of the aircraft is improved. In addition, the take-off weight of the aircraft is reduced, the take-off and landing distance can be shortened, the effective load of the aircraft is improved, and therefore the overall flight performance of the aircraft is improved.

The present invention relates to a method for analyzing tailspin characteristics. The time history of tailspin characteristics is obtained by solving tailspin dynamics differential equations, and the tailspin characteristics of simulation examples under conditions of multi-parameter changes such as weight, moment of inertia, and entry mode are predicted, and the influence of multi-variable parameters on the tailspin characteristics is analyzed; a machine learning model is built using an XGBoost regression algorithm based on decision tree analysis, and the multi-variable parameters of the simulation example are analyzed for feature importance with tailspin characteristics such as average angle of attack, average rotation rate, and sinking speed as targets; the influence ratio of multi-variable parameters on the same tailspin characteristics is analyzed, and the influence of key factors on the tailspin is completed in a quantitative manner. The tailspin feature analysis method of the present invention takes tailspin characteristics as targets, performs feature importance analysis on three multi-variable parameters of weight, moment of inertia, and entry mode, gives the degree of influence of each variable parameter on the corresponding target in the form of quantitative weights, and completes the multi-dimensional global feature analysis of the simulation example.

The invention discloses a device and method for achieving vertical take-off and landing of an aircraft based on control over rotation of a universal joint. The device and method are suitable for switching of the aircraft among a vertical take-off mode, a horizontal flight mode and a vertical landing mode. The device comprises an aircraft body, at least four groups of power devices and universal joint mechanisms matched with the power devices, and the air injection directions of the power devices are adjusted by controlling the universal joint mechanisms to rotate, so that the thrust requirements of different flight modes are met. The control method comprises the steps that real-time state parameters are obtained through flight mode detection, and a vertical take-off mode is initially set; when the aircraft reaches a preset height, gradually adjusting the air injection direction to be horizontal, and entering a level flight mode; and when the landing instruction or the height is lower than a preset value, entering a vertical landing mode, and adjusting the air injection direction to be downward so as to realize stable landing. The operation flexibility and stability of the aircraft are improved, the aircraft is suitable for taking off and landing in a narrow or complex environment, and the energy utilization efficiency is remarkably optimized.

This invention belongs to the field of variable wing technology and discloses a lightweight integrated aerodynamic drive system and a deformable wing. The aerodynamic drive system includes: a double-acting cylinder integrated with the valve bodies of a normally open and normally closed electromagnetic directional valve via a cylinder head; a combined valve comprising an inflation valve and a pressure reducing valve; the inflation valve is located above the pressure reducing valve, with its valve body integrated with the pressure reducing valve's valve body, and its outlet flow channel directly connected to the pressure reducing valve's inlet flow channel; the pressure reducing valve is located on the top of the front cylinder head of the double-acting cylinder, and is connected to the normally open electromagnetic directional valve via a gas-sealed connection, directly docking with a gas cylinder. This system combines optimized gas path design and hardware integration, integrating energy and drive functions, resulting in a compact and lightweight structure. It can adapt well to the requirements of limited installation space, significantly improving the system's adaptability to various operating conditions, thereby enhancing the aircraft's flight performance.

The present invention provides a control and deployment locking system for a tube-launched UAV Π-shaped folding wing, and belongs to the technical field of tube-launched UAV control. In addition to a rotation drive module for driving the main wing and the outer wings on both sides to rotate and deploy from a folded position and a deployment locking module for locking them to a fully deployed position, the system is also provided with an outer wing folding locking module for locking the outer wings on both sides to a folded position and a control module for controlling the time-sharing unlocking of the outer wings, so that the wing on one side rotating forward unlocks the outer wing when the main wing rotates to be perpendicular to the fuselage, and the wing on one side rotating backward unlocks the outer wing before the other outer wing at a predetermined time, generating a rolling restoring torque to balance the rolling torque generated by the difference in incoming flow speed on the wing surfaces on both sides at the initial stage of the main wing deployment, solving the problem that the control mode in which the outer wings are folded in the UAV launch tube by relying on the fuselage constraint and are deployed autonomously after the UAV is launched from the tube causes a large deviation in the rolling attitude of the UAV after it leaves the tube.

The invention relates to an airfoil for a lifting element of an aircraft. The airfoil includes upper and lower contours formed by convex curves and intersection points of the convex curves, the upper and lower contours having given coordinates relative to a chord, the airfoil having a rounded leading edge. For the given chord length and maximum relative thickness of the airfoil, the relative coordinates # imgabs1 # # imgabs2 # of the convex curve of the airfoil are calculated from the midline of the airfoil at a position up to a relative distance # imgabs0 # from the leading edge. The maximum relative thickness of the airfoil is 10% to 25%. The airfoil of a lifting element of an aircraft allows a tail tab to be mounted behind the trailing edge of the airfoil, the tail tab having a rectangular or trapezoidal cross-sectional shape, including a curvilinear shape. The minimum resistance of the airfoil is reduced, the value of the maximum aerodynamic mass is increased, and a smaller value of the pitch moment at zero lift is provided.

The application belongs to the field of insect breeding, and specifically discloses application of citric acid or capric acid in improving flight ability of irradiated sterile male codling moth, wherein the codling moth is fed with a sucrose solution containing 4 mM citric acid or a sucrose solution containing 250 mu M capric acid within 24 hours after the irradiated male codling moth emerges from the cocoon. Compared with the prior art, the application proves that the feeding treatment of citric acid or capric acid can improve the flight ability of the irradiated sterile male codling moth, and further improve the mating success rate of the sterile male codling moth and the control effect on the population.

The invention discloses a fixed time increment fault-tolerant control method for a control surface mixed fault aircraft. The method comprises the following steps: analyzing the fault type and performance of an aircraft control surface; modeling a control surface fault aircraft; improving the design of a tracking differentiator; designing an incremental dynamic inverse control law; and designing a fixed time increment dynamic inverse fault-tolerant controller. The control method provided by the embodiment of the invention is low in cost and simple in structure, can realize accurate acquisition of the angular acceleration signal and noise filtering, reduces hardware consumption of the angular acceleration sensor, and has a practical engineering application prospect; the robustness is high, and stable control and instruction tracking of the aircraft under the control surface mixed fault can be achieved; the method is high in convergence speed and good in steady-state performance, can guarantee that the faulty aircraft quickly recovers to be stable in a fixed time and accurately tracks the instruction, and effectively improves the flight performance of the faulty aircraft.

The utility model relates to the technical field of communication protection, in particular to a small pod device for unmanned aerial vehicle-mounted radar, which comprises a pod outer shell structure, the pod outer shell structure comprises a pod body, a pod top cover, hatch covers and a pod door, the pod top cover is arranged at the top of the pod body, the hatch covers are arranged on two sides of the pod body, and the pod door is arranged in the pod body. The pod doors are arranged on the front face and the back face of the pod body, a round hole clamping groove is formed in the top of the pod top cover, outer heat dissipation holes are reserved in the two pod covers, the inner shell structure is arranged in the pod outer shell structure, and the inner shell structure comprises a battery storage cabin, a host storage cabin and an airborne radar storage cabin; compared with a small pod device of a traditional unmanned aerial vehicle-mounted radar, the small pod device of the unmanned aerial vehicle-mounted radar adopts high-performance light pure aluminum materials and a compact structural design, miniaturization and light weight of the pod device are achieved, the weight and resistance of the unmanned aerial vehicle are reduced, and the flight performance and the cruising ability are improved.

This application provides a compound variable-pitch propeller system and control method based on circulation control, relating to the field of aviation technology. The compound variable-pitch propeller system based on circulation control includes: a dual-rotor variable-pitch motor, which is disposed in the propeller hub. The dual-rotor variable-pitch motor includes a first rotor and a second rotor. The first rotor is connected to an inner worm gear, which is connected to a variable-pitch mechanism. The second rotor is connected to an outer worm gear, which is connected to a control surface deflection mechanism. The propeller blades include fixed blades and movable control surfaces. The fixed blades are connected to the propeller hub, and the movable control surfaces are rotatably connected to the fixed blades. The propeller blades have jet slits. A transmission mechanism connects the control surface deflection mechanism and the movable control surfaces for the deflection motion of the movable control surfaces, thereby improving the efficiency and adaptability of the propeller and optimizing flight performance.